Aircraft fuel storage and distribution system



Marn `l1, T969 E. w. DELANEY 3,432,121

AIRCRAFT FUEL STORAGE AND DISTRIBUTION SYSTEM Filed April s, 1967 sheet l nr e March 115751969 E. w. DELANEY 3,432,121

AIRCRAFT FUEL STORAGE AND DISTRIBUTION SYSTEM Filed April s, 1967 Sheet 2 of e Fl G. 2 50\.

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March 1l, Ef w.' DELANEY AIRCRAFT FUEL STORAGE AND DISTRIBUTION SYSTEM sheet 4- nf@ Filed April 5, 1967 March 1l, 1969 E. w. DELANEY 3,432,121

AIRCRAFT FUEL STORAGE AND DISTRIBUTION SYSTEM Filed April 3. 1967 Sheet 5 of 6 F'IG-6 i W di March ll, 1969 E. w; DELANEY AIRCRAFT FUEL STORAGE AND# DISTRIBUTION SYSTEM sheet" efe Filed April 3. 1967 FIGJO 3,432,121 AIRCRAFT FUEL STORAGE AND DISTRIBUTION SYSTEM Everett W. Delaney, Stratford, Conn., assignor to United Aircraft Corporation, East Hartford, Conn., a corporation of Delaware Filed Apr. U.S. Cl. 244-17.11 Int. Cl.A B64c, 27/04; B64d 37/00 3, 1967, Ser. No. 628,108

27 Claims ABSTRACT OF THE DISCLOSURE An aircraft fuel storage and distribution system with provisions for in-iiight pressure refueling, ground pressure refueling, ground gravity refueling, fuel transfer between tanks, jettisoning of auxiliary tanks, rapid overboard dumping of fuel fr-om all tanks, and selective fuel shut-olf of the fuel being pumped into each tank.

The invention herein described was made in the course of or under a contract or subcontract thereunder with the Department of the Navy.

Background of the invention This invention relates to aircraft, and more particularly to a helicopter which has capability for in-ilight pressure refueling, ground pressure refueling and ground gravity refueling.

In the aircraft refueling art, iixed wing aircraft have been refueled in flight, however, there are particular problems encountered with respect to in-flight refueling of a helicopter due to its rotating lift rotor above the fuselage. It is important that the refueling apparatus be such that the blades of the helicopter lift rotor do not contact any of the refueling apparatus to avoid damage to the parts.

Summary of the invention The primary object of the present invention is to provide a fuel receiving storage and distribution system for a helicopter which is capable of in-flight pressure refueling, ground pressure refueling and ground gravity refueling.

It is still a further object of this invention to utilize the compressor air from the gas turbine engines which are used to power the helicopter lift rotor to both actuate the in-ilight presure refueling probe `and to pressurize the fuel tanks selectively to cause fuel transfer between the tanks as desired.

It is still a further object of this invention to use auxiliary fuel tanks which can be jettisoned and which have fuel and air-pressure connections which automatically shut off in response to the jettisoning action of the auxiliary fuel tanks.

It is still a further object of this invention to teach such a fuel system in which the fuel entering the tanks can be either automatically controlled or controlled by the pilot so that a selected amount of fuel only enters each tank.

It is still a further object of this invention to utilize an interior auxiliary tank which is capable of being selectively gravity fed to either of the main tanks and also be gravity fed to the overboard dump line.

It is still a further object of this invention to teach such a helicopter fuel system which provides for venting of all fuel tanks during refueling and for rapid overboard dumping of the fuel from any of the tanks selectively.

It is still a further object of this invention to teach such a helicopter fuel system having auxiliary fuel tanks which are jettisonable and having tank ll means, fuel and air shut-olf means, fuel transfer means, and tank and system vent means all of which are a part of the airframe and do not drop when the tanks are jettisoned.

ICC

It is still a further object of this invention to teach such a helicopter fuel system which signals the pilot when and which tanks are being filled, when tank filling stops, and Vwhich permits isolation of individual tanks to permit action for damage control in the event of system damage.

Brief description of the drawings FIG. 1 is a showing of a modern helicopter utilizing this invention.

FIG. 2 is a schematic showing of the fuel storage and distribution system of such a helicopter.

FIG. 3 is across-sectional showing Iof the in-flight refueling probe, its actuator and its actuator control system.

FIG. 4 is a schematic showing, including an electrical diagram of the in-flight probe actuation system and its automatic sequencing mechanism.

FIG. 5 is a showing partially in cross section, of the shut-off valve and the high-level sensor fused in the main fuel tanks of this fuel system.

FIG. 6 is a showing, partially in cross section, of the air and fuel drop valve combination used with the jettisonable auxiliary fuel tanks in this fuel system.

FIG. 7 is a view taken along line 7 7 of FIG. 6 to show the bypass line with its particular configuration and thermistor arrangement.

FIG. 8 shows the jettison attachment between the helicopter sponson and one of the auxiliary fuel tanks.

FIG. 9 illustrates the release mechanism which may be actuated by the pilot to jettison either of the auxiliary fuel tanks.

FIG. 10 is a showing of the auxiliary fuel tanks attached to the sponsors of the helicopter.

Description of the preferred embodiment The invention will now be described in particularity by describing the drawings in detail.

Referring to FIG. 1 we see modern helicopter 10 which includes fuselage 12` supporting lift rotorl 14 for rotation and which also includes anti-torque tail rotor 16. Sponsons such as 18 project laterally from opposite sides of the helicopter fuselage 12. The fuselage is supported above the ground by landing gear 20 and 22 which are preferably of the tricycle type. One or more jet-type engines 24 which consist of an air compressor section, a burner section and a turbine section in conventional fashion is mounted in fuselage 12 and provides the power to drive rotor 14. Lift rotor 14 includes at least two blades such as 26 which rotate about rotor head 28 to define a rotor disc having a periphery indicated at 29 and its forward end. Helicopter 10 may be of the type described in greater particularity in U.S. Patents Nos. 3,105,659, 3,097,701 or 3,080,927.

In-flight pressure refueling probe 30 is supported from fuselage 12 by one or more detachable clamps 23, and

. in a fashion to be described hereinafter, is actuatable between a retracted position and an extended position. In FIG. 1, the refueling probe is shown in its extended position and it is important that probe 30, when in its extended position, projects beyond the yforward periphery 29 of the disc defined by the rotating rotor blades 26. The purpose of this is to insure that the apparatus which extends from the refueling plane and which engages inflight refueling probe 30, such as a drogue, does not come in contact with the rotating blades 26 of the rhelicopter.

Referring to FIG. 2 we see a schematic of the fuel system taught herein.` It is important to note that this fuel system is capable of being refueled in three different fashions. First, the fuel system is capable of being refueled by pressure refueling on the ground, secondly, the system is capalble of being refueled by gravity refueling on the groun-d and finally, the system is capable of being refueled by in-tlight pressure refueling.

Ground pressure refueling During ground pressure refueling, a pressurized fuel supply is provided either by some stationary ground installation or by a fuel truck and is connected to ground pressure refueling inlet or adapter 32. This pressurized fuel then ows through line 34 to cross connector 36. This pressurized fuel is prevented from owing to inight probe 30 through line 38 by check valve i401. Check valve 40, and the other check valves referred to hereinafter are of conventional type for example, of the type shown in Perrys Chemical Engineers Handbook, Fourth Edition, pages 6-38 and 6-39, published by McGraw- Hill, which permit flow through a given conduit in one direction only. From cross connector 36, this pressurized fuel may pass through ow sensor 42 and line 44 into the lforward main fuel tank 46. Flow sensor 42 and the other liow sensors described hereinafter, may be of the type disclosed in Bulletin 1070 published by the Revere Corporation of America, Wallingford, Conn. and entitled Flow Switches and serve to provide a signal to the pilot when fuel is flowing therethrough. Each such iiow sensor is connected to the pilots compartment electrically through a conduit such as 48 and provides an indication such as illuminating electric light 541 in the pilots compartment. In entering forward tank 46, the fuel passes through shut-off valve 52. High level sensor 54 serves to automatically actuate shut-off valve 52 when the fuel entering forward tank 46 reaches a preselected level. Also the pilot may manually actuate high level sensor 54 to actuate shut-off valve 52 at the time and level selected by the pilot. The operation of shut-olf valve 52 and the high level selector 54 will be described in greater particularity hereinafter. Surge relief valve 60 is positioned between the shut-off valve 52 and the high level selector 54 and serves the function of preventing damaging pressure buildup in the fluid system. This pressure build-up is caused by both fuel pressure and fuel momentum in the high lrnass flow system. Surge relief valve 60 and the other surge relief valves discussed herein may be of the type shown in U.S. Patent No. 2,396,233 and in the publication entitled Circle Seal Valves published by the Circle Seal Products Company, Inc., 2181 E. Foothill Boulevard, Pasadena, Calif. and identified as 500 series relief valves therein. Continuing our consideration of the pressure re-y fueling operation on the ground, the pressurized fuel from cross connector 36 may also ow through line 64 to cross connector 66, and may pass therefrom either into aft main fuel tank 68, port auxiliary fuel tank 70 or starboard auxiliary fuel tank 72.

The pressurized fuel passing from cross connector 66 to the aft fuel tank 68 passes through `flow sensor 76 in line 78 and eventually through shut-off valve 80 and thence into aft tank 68. Conduit 78 has surge relief valve 81 therein for the purpose of preventing part rupturing pressure build-up therein. High level sensor 82 in aft tank 68 serves the same function as its counterpart 54 in forward tank 46. That is, to actuate shut-olf valve 80 when the fuel entering aft tank 68 reaches a preselected level. Surge relief valve 84 in aft tank 68 serves the same function as its counterpart 60 in the forward main fuel tank 46.

Fuel from cross connector 66 may flow into port auxiliary tank 70` when pilot actuated motor operated gate valve 90 is open in conduit 94. Pilot operated gate valve 90 may be of the type described more fully in the publication of the General Controls Co., 801 Allen Avenue, Glendale, Calif., and entitled, Fuel, Hydraulic and Pneumatic Fluid Controls for Aircraft, Missiles and Ground Support. Flow sensor 916 is also located in conduit 94 to advise the pilot when -fuel liow is occurring through conduit 94, Pressurigd fuel yfrom conduit 94 enters port auxiliary tank 70 through drop valve assemlbly 98.

Pressurized fuel from cross connector 66 may also pass into starboard auxiliary tank 72 through conduit 100 when pilot operated gate valve 102, which is similar to pilot operated gate valve 90, is actuated. Conduit also incl-udes a flow sensor similar to ow sensor 42. This flow sensor is designated as 104. The pressurized fuel in conduit 100 enters starboard auxiliary tank 72 through drop valve assembly 110.

Ground gravity refueling The -main tanks 46 and 68 may be gravity refueled on the ground through ports 112 and 114. Ports 112 and 114 incorporate valves which serve as pressure relief valves for these main fuel tanks 46 and 68 during all types of refueling. Fill-port-relief valves 112 and 114 may be of the type shown in U.S. Patent No. 2,396,233.

In-flight pressure refueling During in-tlight refueling, refueling probe 30 must first be extended. Probe 30 is lactuated by the compressed air from either or both of the compressor sections and 122 of gas turbine engines 24a and 24b. While two such gas turbine engines are shown, it will be understood that the number of engines is immaterial. For purposes of a description we will stress the system as it relates to port engine 24a only. Gas turbine engines 24a and 2417 may be of the type shown in U.S. Patents Nos. 2,711,631 and 2,747,367.

To extend the probe 30, the pilot first opens probe air gate valve 124 located in conduit 126. Air restrictor 128, which is of conventional venturi design, connects conduit 126 to compressor section 120 of gas turbine engine 24u. Check valve 131 insures one-way flow of air through conduit 126. Air restrictor 128 prevents excessive loss of compressor air if there should be a rupture in conduit 126 or in any of the other conduits to which conduit 126 attaches. The compressor section air passing through conduit 126 is lfiltered as it passes through filter 133 and eventually Iarrives at pilot operated selector valve 134, described later in connection with FIG. 3. Selector valve 134 serves two functions. It ports the engine compressor air to one side of the piston-cylinder actuator 136 of in-flight probe 30` through either line 138 or 140 and vents the opposite side of the cylinder-piston arrangement 136 to vent line 142, which is connected to the remainder of the fuel vent system 144. The mechanism for extending and retracting probe 30 will be described in greater particularity hereinafter.

Vent system 144 includes lines 939, 142, 145, 147, 149, 151, 153, 171, 169, 175, 173 and 155 and vents all fuel tanks. Valves 135, 139, and 137 permit selective tank venting.

With in-flight pressure refueling probe 30 so extended and with the probe attached to the appropriate mechanism, such as a drogue, from the refueling aircraft, fuel under high pressure enters through probe 30, past check valve 40 and through conduit 38 to cross connector 36. From cross connector 36 the pressurized fuel may flow to either the forward tank 46, the aft tank 68, the port auxiliary tank 70, the starboard auxiliary tank 72 precisely in the same fashion as described above in connection wit-h the ground pressurized fueling system.

Fuel transfer between tanks In this fuel system, fuel may be transferred between auxiliary tanks and from either .auxiliary tank to either main tank.

Fuel may be transferred from the starboard auxiliary tank 72 to the port auxiliary tank 70 by opening the pilot operated gate valves 90 and 102 and then opening gate valve 130 in pressure conduit 132 and by also opening vent valve in vent system 144. At the same time, vent valve 139 is kept closed and gate valve 137 in pressure line 143 is also kept closed. With these valves so positioned pressurized air from the compressor selection 120 of jet type engine 24a, which is prevented from passing into conduit 126 by the closed condition of valve 124, flows through gate valve 130 and conduit 132 into starboard auxiliary tank 72 to force its fuel therefrom through drop valve assembly for transmission through conduits 100 and `94 into port auxiliary tank 70. It will therefore be seen that by the use of pressurized air from the compressor section of turbine-type engine 24a, fuel may be pumped from the starboard auxiliary tank 72 into the port auxiliary tank 70.

In like fashion, by closing valves and 135 and opening valves 139 and 137, the pressurized air from engine 24a may be used to pump the fuel from the port auxiliary tank 70` to the starboard auxiliary tank 72.

With the valves positioned as just described, it is also possible to use compressor air to force the fuel from the port auxiliary tank 70 into either of the main tanks 46 or 68. This is done when gate valve 90 is open but gate valve 102 is closed so as to prevent flow of this fuel into the starboard auxiliary tank 72. With the valves so positioned, the fuel from port auxiliary tank 70 will ow into both forward main tank 46 and aft main tank 68 through their respective shut-off valves 52 and 80 as previously described in connection with the ground pressure refueling. If it is desired to transfer the fuel from the port auxiliary tank 70` to the forward main tank y46 only, this may be accomplished by pilot actuation of the high-level sensor l82 of aft tank 68 so that the tank shut-off valve 180 will close and accordingly all fuel passing from the port auxiliary tank 70 will pass into forward tank 46 only. Conversely, with the shut-off valve 52 of the forward main tank 46 closed and the shut-oil? valve '80` of the aft main tank 68 open, the fuel from the port auxiliary tank will flow to the aft tank 68 only.

By proper valve positioning it is also possible to use the compressor air from engine 24a to transfer the fuel from starboard auxiliary tank 72 to either the forward main tank 46 or the aft main tank 68. For example, with valves 90, 139 and 137 closed and with valves 130 and 102 and open, compressor air from engine 24a will force the fuel from the starboard auxiliary tank 72 through conduit 100 and then through conduits 64 and 78 to forward tank 46 and aft tank 68. Again, if it is desired for the fuel from the starboard auxiliary tank 72 to go to one of the main tanks only, the pilot may actuate the high level -sensor of the other main tank so as to close the shut-off valve thereof.

While substantial air pressure is necessary to actuate in-ight pressure `refueling probe 30, this substantial amount of air pressure is not necessary to cause fuel to transfer from the auxiliary tanks and may well damage the tanks or system, It is therefor desirable to regulate the pressure of the air being passed from the engines through conduits 132 and 143 to the auxiliary tanks 70 and 72. Accordingly, pressure regulators 940 and 942 are placed in conduits and 132, respectively. Pressure regulators 940 and 942 serve to regulate the pressure of the air being passed from the engines 24a and 24b to the auxiliary tanks 70 and 72 to prevent tank rupture. These pressure regulators may be of any variety, for example, of the type shown in Perrys Chemical Engineers Handbook, Second Edition, published by McGraw-Hill and illustrated on pages 2032 through 2035 thereof. It would also be possible to substitute an integral unit for each of valve 130 and pressure regulator 942, and valve 137 and pressure regulator 940 by using a valve-pressure regulator combination mechanism of the type described in the publication of the Janitrol Aircraft Division of Midland- Ross Corporation, 4200 Surface Road, Columbus 4, Ohio entited Valve Control Assembly-Air Pressure and identited as Data Sheet JA-128.

Rapid fuel dumping With this fuel system it is possible to dump fuel from either of the auxiliary tanks 70 and 72 or the main tanks 46 :and 68 overboard rapidly. This can. be accomplished by forcing the fuel from the auxiliary tank into either of the main tanks by the use of compressor air in the fashion described above and then utilizing rapid dumping fuel pump which is of the type more fully described in Perrys Chemical Engineers Handbook, Second Edition, published by McGraw-Hill, pages 22452270, to pump the fuel from the main tanks 46 and 68 through conduits 152 and 154, respectively, and then through conduit 156, with check valve 158 therein, overboard through overboard line 170. Pilot operated gate valves 172 and 174 are located in conduits 152 and 154, respectively, and' serve to connect either forward main tank 46 or aft main tank 68 or both such tanks with pump 150. If the fuel in aft tank 68 only is to be pumped overboard, valve 172 is closed while valve 174 is opened and the action of pump 150 will serve to pump the fuel from tank 68 only overboard. Conversely, if the fuel from forward tank 46 only is to be pumped overboard, valve 172 is open and valve 174 is closed. It is also possible to pump the fuel from both tanks 46 and 68 overboard rapidly by opening both valves 172 and 174. In addition, by forcing the fuel from either port auxiliary tank 70 or starboard auxiliary tank 72 into one or both of the main tanks` 46 and 68, the fuel from the auxiliary tanks may also be pumped overboard rapidly from the main tanks by the action of pump 150.

Each of the main tanks 46 and 68 and the auxiliary tanks 70 and 72 are attached to the fuel systems vent system 144. Vent line 169 and vent line 171 are connected to tanks 68 and 46, respectively. Vent lines 173 and 175 also extend from these main tanks 68 and 46. Vent valve 135 can be actuated to connect the port auxiliary tank 70 to either vent line 169 or 171 while the vent valve 139 can be pilot actuated to connect the starboard auxiliary tank 72 to either the vent line 173 or 17 5.

Still referring to FIG. 2 we see interior auxiliary tank which is mounted in fuselage 12 to be higher than either forward tank 46, aft tank 68 and overboard dump line 170 so that the fuel from this tank may be gravity transferred to either the forward tank 46 through conduit 182 when pilot operated gate valve 184 is open or to aft tank 68 through conduit 186 when pilot operated gate valve 188 is open. Additionally, when pilot operated gate valve 190 is open, the fuel from interior auxiliary tank 180 may be gravity drained overboard through conduit 192 which connects to overboard drain line 170.

As best shown in FIG. 2, it will be noted that probe 30 and its actuating system 136 are self-contained and may be quickly attached or detached from the fuselage because of probe clamps 23 and quick attach-detach couplings 800, 802, 804 and 806 in air lines 138, 140 and 139 and in fuel conduit 38. These couplings may be of the type described in Perrys Chemical Engineers Handbook, Second Edition, published by McGraw-Hill, page 925.

Referring to FIG. 3 we see my extensible-retractable in-flight pressure refueling probe 30 in greater particularity. The probe actuating mechanism 136, which i-s of circular cross section and concentric about axis 254, consists of a stationary cylinder assembly 250 within which movable piston-probe mechanism 252 can reciprocate. Cylinder assembly 250 includes outer cylinder 251 and inner cylinder 253 which are sealably joined by end plates 25S and 257 to define annular chamber 259 therewithin. Both cylinder assembly 250 and piston-probe mechanism 252 are preferably of circular cross section and concentric about axis 254 and made of aluminum for the purpose of obtaining maximum weight-to-strength ratio. While the double-walled cylinder provides a rugged construction and the necessity for fuel conduit 253 lends itself to this construction, it will be recognized that outer cylinder 251 and piston-probe 252 define pressure chambers 282 and 284. Piston-probe mechanism 252 consists of outer cylinder member 256 and inner cylinder member 258 joined thereto by piston 260 and member 261 for strength purposes. Piston member or skirt mechanism 260 extends radially outwardly from cylindrical members 256 and 258 and coacts with the inner surface of cylinder assembly 250, through the cooperation of circumferential seal 262, to form a cylinder-piston arrangement therewith and define pressure chambers 282 and 284 therewithin on opposite sides of piston 260. Circumferential seals 263 and 265 extend and seal between cylinder assembly 250 and piston-probe assembly 252. As best shown in FIG. 3, the sliding surfaces such as surfaces 818, 812 and 814 between piston assembly 252 and cylinder assembly 258 are preferably hard anodized by a conventional hard anodizing process, for example, of the type described in U.S. Patents Nos. 2,692,851 and 2,692,852. Inner cylindrical member 258 extends forward beyond outer cylindrical member 256 and attaches to and supports probe tip 264. Members 258 and 253 cooperate to define probe fuel conduit 27 9 which connects to conduit 38.

By viewing FIG. 3 it will be noted that the fuel which enters probe 30 passes both through the probe and its actuating system through internal conduit 279 defined by cylinder members 258 and 253, thereby avoiding the drag creating problems encountered with fuel conduits which extend external of the probe. Also, because these two conduits are always aligned with fuel conduit 38, it will be noted that probe 30 is ready for refueling when in any position between its retracted and extended positions Forming an integral part of the cylinder assembly 250 are the retracte abutment ring 811 and the extend abutment ring 813 which serve as stops to limit the throw of the piston-probe assembly 252 within the cylinder assembly 250. Elements 811 and 813 are shown as rings but could be radial pins which are retractable to permit piston removal. Piston-probe assembly 252 contains two 'axially spaced circumferential grooves 266 and 268 on opposite sides of seal 262, which cooperate with lock mechanisms 270 and 272 to lock piston-probe assembly 252 in its retracted position shown in solid lines in FIG. 3 and in its extended position shown in phantom in FIG. 3. Lock mechanisms 270 and 272 are identical and preferably comprise electric motor 275 and 274 which causes lock pin 277 and 276 to rotate or otherwise move radially inwardly to engage annular recesses 266 or 268 to lock the probe-cylinder mechanism 252 in either its extended or retracted positions. Lock pin 276 of motor 272 engages recess 266 for locking probe-piston assembly 252 in its extended position while lock pin 277 of motor 270 engages recess 268.

As best shown in FIG. 3, air from the compressor section 120 or 122 of one or more of the engines 24 passes through line 126 and then through open valve 124 and filter 133 to selector valve 134. This compressed air passes through conduit or port 280 in multi-positionable selector valve 134 and then either through conduit 138 to 140 to the opposite pressure chambers 282 and 284, on opposite sides of piston member 260 of probe-piston 252 to cause .the probe to either extend or retract. It will be noted that selector valve 134 includes recesses 298 and 292, which serve to vent the side of probe-piston mechanism 252 which is not under pressure. As shown in FIG. 3, recesses 290 and 292 vent pressure chambers 282 and 284 when the selector valve 134 is in its neutral position, thereby removing the motive force of the compressed air from probe-piston assembly 252 and the attendant pressure variations which would otherwise be encountered therein as the altitude of the helicopter changed. In operation, selector valve 134 moves from the neutral position from which it is shown in FIG. 3 to a rst position in which conduit 280 communicates with conduit 140 to provide compressor air into chamber 282 so as to propel probepiston mechanism 252 to its extended position, during which selector valve position, recess 290 puts conduit 138 and hence chamber 284 into communication with vent conduit 142. Vent line 939 with check valve 141 therein connects fuel conduit 38 to vent line 142 to prevent a conduit collapsing vacuum from forming in line 38 between probe 30 and check valve 40 when the probe is being extended. When the refueling probe 30 is to be retracted, the selector valve 134 is rotated so that the converse occurs, namely, that compressed air is supplied to chamber 284 through the coaction of conduits 138 and 280 while chamber 282 is vented to the vent line 142 by the coaction of conduit 140 and recess 292. Check valve 141 performs the function of preventing fuel from conduit 38 from passing into the vent lines 939 and 142. Limit switches 300 and 302 are actuated by piston member 260 to provide a signal when the probe 30 is at the opposite ends of its travel so as to bring about certain operations in the sequencing of the probe now to be described.

Still referring to FIG. 3 it will be seen that in-flight pressure refueling probe 30 and its actuating System are electrically bonded to the fuselage 12 of aircraft 10 and the parts of probe 30 and actuator 136 are electrically bonded because probe-piston member 252 is electrically connected to outer cylinder 251 at all times by leaf spring member 500 and outer cylinder 251 is electrically connected to fuselage 12 through clamps 23. Spring member 500 attaches to member 252 and bears against the inner wall of outer cylinder 251 for all positions of member 252. This electrical bonding between the entire probe unit 30 and the helicopter 10 insures against electrical discharge between the refueling aircraft and the helicopter and the fire danger created thereby.

While our preferred embodiment uses a fixed cylinder 250 and movable piston 252, it is within the Skill of the art to make piston 260 fixed and to move probe tip 264 with cylinder 250.

Sequence of operations for probe extension and retraction The sequence of operations of probe 30 is best described with respect to FIG. 4. For the purposes of explanation, the probe 30 and mechanical locks 270 and 272 are schematically supreimposed on the electrical circuitry which controls be prolbe movements. It should be noted that all relays in FIG. 4 |are shown in the deactived state.

Assuming that the probe is retracted and locked, the following sequence takes place. When the pilot desires to execute a refueling mission, he must first turn on the tmaster switch 700 to energize his selector switch 702, gate valve 124 for compressor bleed air and other appropriate circuitry from the 28 v. DC power brus.

Extension of the probe The pilot throws his selector switch 702 to the Extend position. Relay 704 is energized to establish the proper polarity across lock pin actuator 275 for retractin-g the lock pin 277. As the actuator 275 retracts the lock pin 277, microswitch 706 will be thrown immediately after the lock pin 277 begins to be withdrawn from the locking recess 268. As microswitch 706 is thrown, the probe-not stowed light 708 is energized to indicate to the pilot that the refueling probe 30 is no longer locked. As the actuator 275 moves the lock pin 277 to its fully disengaged position, microswitch 710 will be thrown to shut off electrical power to the actuator 275 and to energize selector valve 134 through the contacts of relay 712. Relay -712 is activated by the 28 v. DC power bus whenever the other lock pin 276 is Withdrawn. The selector valve 134, being a three-positioned valve, will be moved from its deenergized neutral position to extend the probe by applying compressor lair from the gate valve 124 to the extend chamber 282 and propel the piston assembly 252 toward the extend probe position. As the probe begins to extend, limit switch 300 will be thrown and electrical power for the actuator 275 will tbe completely removed from the contacts of relay 704.

The probe 30 will move to its extended position and limit switch 302 will be thrown by the piston assembly 252. Relay 714 will be actuated through the contacts of microswitch 710 and limit switch 302 to establish the proper polarity across look pin actuator 274 `for engaging lock pin 276 with the locking recess 266. The actuator 274 will drive lock pin 276 into the piston assembly 252 and as the lock pin 276 fully engages the recess 266, microswitch 716 will be thrown to deenergize the actuator 274, relay 712 and the probe-not-stowed light 708. When relay 712 is ideenergized, selector valve 134 will assume its neutral position and shut olf the bleed air to the piston assembly 252 while venting both chambers 282 and 284. At the same time, the contacts of microswitch 716 also energize the probe-ready light 718 to indicate to the pilot that the probe 30 is locked and prepared to engage the refueling drogue for fuel transfer.

Should mechanical lock 272 fail to lock it will be noted that our system is fail safe because microswitch 716 will not tbe thrown to deenergize relay 712 and selector valve 134 and hence compressor 'bleed air from the selector valve 134 will hold the probe in its extended position until the selector switch 702 is thrown to its Retract position.

Retracton of the probe When the pilot has taken on yfuel and disengaged the probe from the refueling drogue, he retracts the probe by moving his selector switch 702 to the Retract position. Electrical power previously energizing relay 714 through microswitch 710 and limit switch 302 will immediately be lost and the contacts of relay 714 will reverse the polarity previously applied across actuator 27'4. Actuator 274 irnmediately withdraws look pin 276 from the piston lassembly 252. As lock pin 275 begins to move out of engagement with recess 266, microswitch 716 will be thrown to extinguish the probe-ready light 718 and to energize the probe-not-stowed light 708. When the lock pin 276 is completely withdrawn, microswitch 720 will lbe thrown to deenergize actuator 274 and to energize selector valve 134 through the contacts of limit switch 300 and time delay relay 722. The time delay relay 722 is of the type which does not delay the input signal when the relay is pulled in but which produces an artificial signal for a short period f time after the actuating signal has been lost. Consequently, the selector valve 134 will be actuated as soon as the microswitch 720 is thrown and will remain actuated for a brief period after the signal to the time delay relay 722 is lost. The selector valve 134 when actuated by relay 722 moves from its neutral position to apply compressor air to the retract chamber 284 and propel the piston assembly 252 toward the retracted probe position. As the piston assembly 252 moves into the retracted probe position, limit switch 300 will be thrown to electrify the contacts of deactivated relay 704 and to terminate the signal to the time delay relay 722 for the selector valve 1314. The time delay relay 722 holds the selector valve 134 open for an additional period of approximately two seconds to insure that the piston assembly 252 remains in the retracted position While the actuator 275 moves lock pin 277 into engagement with the recess 268. When lock pin 277 is fully engaged with the recess 268, microswitch 706 is thrown to extinguish the probe-not-stowed light 708 and to deenergize the actuator 275i.

It will be noted that the probe-not-stowed light takes electrical power directly from the 28 v. DC power bus and that the light 708 will be energized whenever the probe 30 is not locked even though the master switch 702 is not on. 'Ilhe probe-not-stowed light 708 and the probe-ready light 718 provide the pilot with an indication of the probe condition at all times during the refueling operation 'and at any other time that the probe is not safely stowed and locked.

Referring to FIG. we see a typical main fuel tank, shut-off valve 52 and high-level sensor 54 in greater particularity. While the shut-off valve and the high level sensor of forward main fuel tank 46 will be described, it should be borne in mind that comparable mechanism is present in aft main fuel tank 68. Shut-off valve 52 consists of outer wall member 304 and inner wall member 306 which define annular chamber 308 therebetween, which -annular chamber communicates with the interior of tank 46 through conduit 310. Double area .piston 312 is biased by spring 314 to its closed position shown in FIG. 5. Piston 312 has center conduit 316 passing therethrough and has small area end 318 at one end thereof and large area end 320 at the opposite end thereof. As fuel enters shut-off valve 52 through conduit 44, the pressure of the fuel overcomes the force of spring 314 to force piston member 312 downwardly to permit passage thereby and through ports 322 and 324 into the interior of fuel tank 46. This fuel also passes` through central conduit 316 and into chamber 326 from whence it ows through conduit 328 over to high level sensor member 54. High level sensor member =54 comprises hollow housing member 330 having float member 332 therewithin. Shaft member 334 projects from oat lmember B32 and extends through aperture 336 in housing member 330. A second aperture 338 in the base of housing member 330 permits the fuel within cham-ber 46 to act against iloat member 332. When the fuel in fuel tank 46 is at a level lower than high level sensor 54, float member 332 rests against the base surface 340y of housing member 330 therebly placing conduit 328 into communication with conduit 342 through chamber 344. Conduit 342 is open so that any fuel passing therethrough dumps into the interior of fuel tank 46. As the level of the fuel in tank 46 gets higher, it eventually enters aperture 338 and acts upon float 332 to cause the Ifloat to move higher within chamber 344 until float 332 eventually blocks communication between conduits 328 and 342. The afloat in this condition, the pressure in chamber 326 of the shut-off valve 52 builds up and, due to the difference in area between the ends 318 and 3-20 of piston 312, the piston, with the aid of spring 314, is forced to its closed position, thereby preventing the admission of further fuel into tank 46. It should be noted in FIG. 5 that a solenoid-type mechanism 350 coacts with `shaft member 334 of the high level sensor 54 so that this pilot oper'- ated solenoid member 350 can be actuated to cause float member 3'32 to assume its blocking position, thereby lshutting olf shut-off valve 52 at the will of the pilot.

Referring to FIG. 6 we see starboard drop-valve 110 of the starboard auxiliary fuel tank 72. While drop valve 110 will now be described, it should be borne in mind that the port auxiliary tank 72 has an. identical drop valve 98. In viewing FIG. 6 it will be noted that fuel passes through fuel conduit into chamber 360 de- -lined by housing 362. Housing 362 also houses fuel valve 364 which includes a hollow stem member 366 having peripheral apertures 368 passing through the walls thereof so as to place the hollow stem -member A366 of valve 364 into communication with chamber 360. Spring mem ber 370 biases fuel valve 364 so that the undersurface 372 thereof bears against valve seat 374. As auxiliary fuel tank 72 is placed in position, standpipe 380, projecting therefrom and communicating with the interior thereof, bears against the bottom surface 382 of fuel valve 364 to force fuel valve 364, against the action of spring 370, to its open position so that any fuel which enters chamber 360 through conduit 100 m-ay pass around valve 364 and through apertures 368 into the interior of the hollow stern portion 366 of valve 364 and thence into auxiliary fuel tank 72 through standpipe 380.

Still viewing FIG. 6 it will be noted that a similar arrangement is provided for air passage into and out of auxiliary tank 72 through conduit 132. Air from conduit 132 enters chamber 382 defined by housing 384, which housing also houses air valve 386, which is similar to fuel valve 364 just described, and which yincludes hollow stem member 388 with the peripheral apertures 390 extending through the walls thereof. Spring 392 serves to force air valve 386 to its closed position wherein its lower surface 394 contacts seat 396. Accordingly any air which enters chamber 382 through conduit 132 may pass around air valve 392 whenever that air valve is lifted to its open position by the action of air standpipe 398 which extends from auxiliary fuel tank 72. After passing around valve 386, the air may pass through nozzle-like -apertures 390 into hollow Istem member 388 `of the valve and then through standpipe 398 into the interior of auxiliary fuel tank 72. With air valve 386 in its -open position, any air may also pass from the interior `of auxiliary tank 72 through st-andpipe 398, hollow stem member 388, apertures `390, chamber 382 and be exhausted or vented thr-ough a conduit 132.

Accordingly, in operation, when fuel is being admitted to auxiliary fuel tank 72 through conduit 100 by either the ground pressure refueling system or the in-iiight probe pressue refueling system, the fuel enters chamber 360, passes around valve 364 and through apertures 368 into the interior of hollow stern 366 and thence through standpipe 4380 into auxiliary fuel tank 72. As fuel is being admitted into the auxiliary fuel tank 72, air must be discharged therefrom to permit unimpeded entry of the fuel and accordingly the displaced air is passed through standpipe 398, hollow valve stem member 388, apertures 390, chamber 382 and eventually conduit 132 to vent.

During transfer of fuel from auxiliary tank 72 to either of the main tanks 46 or 68 or the auxiliary tank 70, compressed air from the engine compressor air sections 120 and 122 is forced through conduit 132 and into chamber 382 of the air valve section of drop valve 110 and thence into auxiliary tank 72 through apertures 3901, hollow stern 388 and standpipe 398. This forced entry of regulated air pressure into the auxiliary tank 72 displaces the fuel therein and causes that fuel to be transferred from the tank as it passes through standpipe 380, hollow stern section 366, apertures 368, chamber 360 and conduit 100 to any of the other tanks just mentioned.

When fuel is being transferred from the auxiliary fuel tank 72, vent valve 139 (FIG. 2) is closed so that air pressure will not vent overboard and can be introduced through conduit 132 and the air drop-vent valve 110 into the tank 72. When fuel is being introduced to the auxiliary tank 72 through the fuel drop valve 110, valve 139 is opened to place the outlet of the air drop Valve in communication with vent system 144.

It is an important feature of the drop valve 110` that when the tank 72 is jettisoned, no compressed air or fuel be wasted and therefor that air valve 386 and fuel valve 354 are shut off. This action is caused when the jettisoning of auxiliary tank 72 causes the removal of standpipes 380 and 398 to thereby permit springs 370 and 392 to force fuel shut off and metering valve 364 to air shut off and metering valve 386 to their closed or shut off positions.

lIt is important during the lling operation of the auxiliary fuel tanks 72 that fuel ow to the tank be discontinued as soon as the tank is filled. This is accomplished by the use of thermistors 400 which are electrically connected to the gate val-ve 102 in conduit 100 (see FIG. 2). Thermistors 400' are temperature responsive and serve to sh-ut off gate valve 102 and thereby prevent further flow of fuel to auxiliary tank 72 through conduit 100 when fuel comes into contact with the thermistors 400, thereby indicating that the auxiliary tank 72 is full. It is an important teaching of this construction that the thermistors not be located in chamber 382 of the air dump valve since it has been found that with the thermistors so located, the rapid flow of air from auxiliary tank 72 thro-ugh chamber 382 during refueling causes the highly temperature sensitive thermistors 400 to actuate shut-off valve 102 to its closed position when, in fact, the auxiliary tank 72. is not filled. It is accordingly a teaching of this construction to place thermistors in a particularly designed bypass sampling line, which bypass sampling line 402 communicates at its opposite ends with air dump valve chamber 382. Bypass sampling line 402 is preferabl)l extending at an angle to the vertical direction to permit gravity drainage of liquid therefrom and has at its lower end entrance conduit 404 `which is of smaller crosssectional area than outlet conduit 406. Each of conduits 404 and 406 communicate with plenum chamber 408. Plenum chamber 408 is of substantially larger cross-sectional area than either of conduits 404 or 406 and is defined within walls 410 which, with respect to fluid flowing from 404 to 406, are smoothly divergent at section 412, and smoothly convergent at section 414 and smoothly blended therebetween. The purpose of this particular shape of bypass sampling conduit 402 is to cause a sample of the air being expelled from the auxiliary tank during refueling to pass by thermistors 400 at a low -velocity and in smooth nonturbulent flow so as to have minimum effect upon the temperature sensitive thermistors 400. When during refueling, auxiliary tank 72 is eventually filled, the initial overflow fuel therefrom will pass through apertures 390 of the air drop valve and a portion through bypass sampling passage 402. To insure iiuid iiow through bypass sampling passage 402, it is essential that there be a pressure drop between the inlet and outlet thereof. This pressure drop is caused by positioning inlet 405 of bypass sampling conduit 402 closer to the region of maximum dynamic pressure adjacent nozzle-type aperture 390 than the outlet 407 of conduit 402. Accordingly, due to this pressure drop, initial overtiow fuel from full auxiliary tank 72 passes through bypass sampling conduit 402 and when that fuel first contacts the thermistors 400, those thermistors, due to an electrical connection between thermistors 400 and shut-off valve 102 (FIG. 2) will rapidly close the shut-off or gate valve 10'2 so as to prevent the pumping of additional fuel to the auxiliary tank 72. There is a similar thermistor control in drop valve 98 of auxiliary tank 70 which shuts off valve in conduit 94.

When regulated compressor air passes into auxiliary fuel tank 72 from conduit 132 through the air drop valve to transfer fuel from auxiliary tank 72, such compressed air willralso pass through bypass conduit and clean any fuel deposited therein during refueling. It is also important that thermistors 400 are located at the top of plenum chamber 408 so that fuel will drain by gravity.

There are certain significant features of drop valve and the thermistor system 402 illustrated in the pre- -ferred embodiment shown in FIGS. 6 and 7. Preferably, the area of conduit 132 is larger than the combined area of apertures 390 to insure unrestricted flow through chamber 382 of the air being forced out of auxiliary tank 72 `during refueling. In addition, because the area of conduit 406 is larger than the area of conduit 404 and because of the nozzle pressurizing effect of the fluid (whether air, a mixture of air and fuel, or fuel) passing through apertures 390, the pressure at the inlet 405 to inlet conduit 404 is greater than the pressure at outlet 407 of conduit 406i thereby establishing the necessary pressure drop through bypass sampling conduit 402 to establish fluid flow therethr-ough` The areas of conduits 132 and the total areas of the apertures 390 must be selected so that the pressure drop thereacross is greater than the pressure drop across bypass sampling conduit 402. Also, the area of outlet 407 of conduit 406 must be selected so that it offers less restriction to flow than do apertures 390. This does not always mean that the area of conduit 406 is greater than the combined areas of apertures 390 because there is greater resistance to the fluid passing through the series of apertures 390 than there is to the fluid passing through the single enlarged conduit 406. It should be noted that while apertures 390 are illustrated as being of circular cross section, they could as well be one `or more slots.

As previously described plenum chamber 408 (FIG. 7) performs the dual function of:

(l) reducing the velocity of the uid passing therethrough, (2) producing laminar fluid flow therethrough.

where:

12e-:Reynolds number (dimensionless) p=density of the fluid V=velocity of the uid D=the diameter of the con-duit and, ,w=is the absolute viscosity of the fluid.

A consideration of Reynolds number gives some indication of ow conditions.

With the Reynolds number at about 2,000 or below, laminar flow through plenum chamber 408 is achieved. With the Reynolds number above 2,300 turbulent flow occurs through the plenum chamber and with the Reynolds number therebetween, mixed ow occurs. It is recognized that there is some Reynolds number variation with variation in operating conditions.

Another advantage to be gained by the use of this specifically contoured plenum chamber 408 is -fuel drainage. Actually, during refueling of fuel tank, not only air passes through bypass sampling line 402, but a mixture of air and fuel droplets or foam. The reduced velocity of this mixture as it ows through the plenum chamber causes the fuel to drop due to gravity to the 'bottom of plenum chamber 408 and drain therefrom by the force of gravity. It accordingly might be desirable to provide a removable drain plug (not shown) in the bottom of plenum chamber 408 to remove fuel so accumulated therefrom. This drain plug would be especially useful in installations where pressurized air is not reverse flowed through chamber 382 and bypass sampling line 402 for fuel transfer.

Our experience has shown that the ratio of the dia-meter of outlet 407 to the diameter of inlet 405 of conduit 404 should be no less than 2.5 to 1 for an operation in which fuel enters tank 72 at from 0 to 150 gallons per minute. In addition, due to the fluid flow resistance caused thereby, it would be undesirable to have the diameter of c-onduit 404 less than 1A" in any aircraft fuel system. Referring to FIG. 8 we see the manner in which the auxiliary fuel tanks 70* and 72 are mounted. While port tank 70 only is shown, it will be understood that the star-board tank 72 is mounted in the same fashion, Fuel tank 70 is mounted from sponson 18 or more particularly from pylon 19 extending laterally outwardly from sponson 18 by -means of shackle 600 which is pivotally connected by pins such as 6.02 totapertured lugs 604 and 606 projecting from sponson 18. Shackle 600, which is ibasically hollow, carries therein an actuating system, best shown in FIG. 9, for disengaging pins 608 and 610, which project in pivot fashion through apertured lugs 612 and 614 projecting from auxiliary fuel tank 70'. Pivotal sway braces 620, 622, 624 and 626 are pivotally attached to sponson 18 by pivot pin lugs such as 630 and 632 and include pad members such as 634, 636, 638 and 640 which engage the outer periphery of auxiliary fuel tank 70. In this fashion with shackle 600 pivotally connected to the sponson and to the auxiliary fuel tank 70, and with the sway braces 620 through 626 bearing against the auxiliary tank, the auxiliary tank is held in place but in jettisonable fashion from the helicopter sponson 18. The jettison action will now be described.

Referring to FIG. 9 We see the jettison linkage 690 in enlarged form and extending outside of shackle 600 for purposes of illustration 'but linkage 690 is preferably all housed within shackle 600. The linkage includes pivot links 650 and 652, which are pivotally mounted at xed pivot points 649 and 651 and which `are pivotally connected at one end thereof to pins `608 and 610 at 654 and 656 and are pivotally connected at their other ends at 658 and 660 to cross rods 662 and 664. Cross rods 662 and 664 are pivotally connected to crank members 666 and 668, which are pivotally connected at points 670 and 672 and which are joined by mating ring gears or gear segments 674 and 676 to cause crank pivot motion about these pivot points in unison. One of the cranks such as 668, is pivotally -connected at v678 to actuator arm `680. Actuator army 680 may be caused to move leftwardly either by pilot actuated lanyard attachment 682 or by piston-cylinder hydraulic or pneumatic mechanism 684'. When connector link 680 is caused to move leftwardly as shown in FIG. 9 by either pilot actuated lanyard iconnection 682 or piston-cylinder mechanism 68-4, crank `668 pivots in a counterclockwise direction to cause pivot link 652 to rotate or pivot in a clockwise direction so as to withdraw connecting pin `610 from auxiliary tank projection `614. At the same time, due to the coaction between ring gears 674 and 676, crank 666 vmoves in a clock- Iwise direction to move pivot lever 650 in a counterclockwise direction so as to withdraw connecting pin 608 from pivot lug 612 projecting from the auxiliary tank 70. With both of these lugs removed, they tank 70 will thereby be jettisoned from the aircraft in flight since the sway braces 620 through 626 perform no tank retaining function. It will be evident that with. this type of retension system, the auxiliary tanks can be supported from any surface, whether vertical, horizontal or therebetween.

The in-flight probe 30 and its actuating system illustrated in FIGS. 3 and 4 hereof are being claimed in patent application 628,109 tiled on even date herewith in the name of Joseph DiPiro and entitled, In-Flight Pressure Refueling Probe and Actuation System. The auxiliary tank shut-off valve and its overflow sensing system illustrated in FIGS. 6 and 7 hereof are being claimed in patent application 628,013 led on even date herewith in the name of Stanley J. Iusyk and entitled, Fluid Tank Shut-Off Valve and Actuation System.

I claim:

1. In `a helicopter having:

(l) a fuselage,

(2) a lift rotor supported for rotation fromI said fuselage,

(3) sponsons projecting laterally from the opposite sides of said fuselage,

(4) at least one gas tunbine engine having an air com- 'pressor section and mounted in said fuselage to drive said lift rotor,

(5) a fuel system including:

(a) at least one main fuel tank located in said fuselage,

(b) `an auxiliary fuel tank supported from each of said sponsons,

(c) an extendable and retractable in-flight pressure` refueling probe actuatable by compressed air from said engine and project-ing forwardly from said fuselage and dimensioned to extend beyond the helicopter rotor when extended,

(d) means connecting said in-ilight probe to said main tank and to said auxiliary tanks to permit inflight pressure refueling thereof,

(e) means connected to said main. tank and said auxiliary tanks to permit ground pressure refueling thereof,

(f) means connected to said main tank to permit ground gravity refueling thereof,

(g) means responsive to the compressor air generated by said air compressor section of said engine to permit fuel transfer from either of said :auxiliary tanks to said main tank,

(h) means to vent all of said tanks: during refuelung,

(i) means operably associated with each of said tanks to determine the amount of fuel to be received in each of said tanks,

(j) means operably associated with each of said tanks to selectively permit rapid overboard dumping of the fuel from each of said tanks.

2. Apparatus according to claim 1 and including quick detach means to support said auxiliary tanks from said sponsons to permit in-flight jettisoning thereof.

3. Apparatus according to claim 2 and including a third auxiliary fuel tank mounted in said fuselage above said main fuel tank, and including means to transfer fuel from said third auxiliary tank to said main fuel tank, and still further including means to rapidly dump fuel yfrom said third auxiliary tank overboard.

4. Apparatus according to claim 1 and wherein said fuel dumping means includes a high volume fuel pump operatively connected to all of said tanks, and including means to select which of said tanks is to have fuel dumped therefrom.

5. Apparatus according to claim 1 and including a second main fuel tank mounted in said fuselage and further including means connecting said second main fuel tank to said in-flight refueling probe for pressure refueling therefrom, means connected to said second main fuel tank for ground pressure refueling thereof, means connected to said second fuel tank for gravity refueling thereof, and means connecting both of said auxiliary fuel tanks to said second fuel tank to permit fuel transfer in response to air pressure from the air compressor section of said engine between either of said auxiliary tanks and said second main tank, and means to dump the fuel from said second main tank overboard rapidly.

6. Apparatus according to claim 3 and including means to prevent the fuel being dumped overboard from said third auxiliary tank from owing into said main tank.

7. Apparatus according to claim 2 and including means responsive to the positioning of said auxiliary fuel tanks into position for support by said sponsons to place each of said fuel tanks into communication with the air compressor section of said engine, said in-flight pressure refueling probe and said ground pressure refueling mechanism, which means is also responsive to the jettisoning of said auxiliary tanks to cut off communication with said air compressor section, said in-flight pressure refueling probe and said ground pressure refueling means.

8. Apparatus according to claim and including surge pressure relief valves communicating with each of said main fuel tanks to prevent rupture thereof due to fuel surge pressure.

9. Apparatus according to claim 1 and wherein said means for determining the amount of fuel to be received by said main tank is pilot operable to permit selection of the amount of fuel to be pumped into said main tank and including a tank shutoff valve responsive to said means to block fuel ow into said tank when the preselected amount has been reached.

10. Apparatus according to claim 1 and including a pressure relief valve in said main tank, which pressure relief valve also serves as the fuel inlet to said main tank during gravity refueling thereof.

11. A helicopter including:

(l) a fuselage,

(2) sponsons projecting laterally from the opposite sides of said fuselage,

(3) a lift rotor supported for rotation from said fuselage,

(4) a gas turbine engine having an air compressor section and mounted in said fuselage to drive said lift rotor,

(5) and a fuel system including:

(a) a forward and after main fuel tank located in said fuselage,

(b) an auxiliary fuel tank attached to and sup- 16 ported from each of said sponsons so as to be jettisonable therefrom in ight,

(c) an extensible and retractable in-ight pressure refueling probe attached to said fuselage and projecting forwardly therefrom and dimensioned so that the end of the probe extends beyond the area of the lift rotor when the probe is extended,

(d) means responsive to the air pressure generated by the compressor section of said engine to actuate said in-flight pressure refueling probe between an extended and retracted position,

(e) means connecting said refueling probe to each of said tanks for refueling each of said tanks through said refueling probe,

(f) a ground pressure refueling adapter supported in said fuselage,

(g) means connecting said ground pressure refueling adapter to each of said tanks to permit refueling of each of said tanks through said ground pressure refueling adapter,

(h) means for determining which of said tanks is to be refueled through either said in-ight pressure refueling probe or said ground pressure refueling adapter.

(i) means for refueling said forward and after main tanks by gravity,

(j) means for shutting off fuel ow to each of said tanks when said tank is lled to a preselected level during refueling,

(k) means to vent said tanks during refueling,

(l) means to transfer fuel from either of said auxiliary tanks to either of said main tanks,

(m) means to transfer between said auxiliary tanks,

(n) and means to rapidly dump fuel overboard from any of said tanks.

12. Apparatus according to claim 11 and further wherein fuel is transferred between said tanks in response to pressurized air from the compressor air section of said engine.

13. Apparatus according to claim 12 and including an interior auxiliary fuel tank mounted in said fuselage above said main fuel tanks, and including means to transfer fuel from said interior auxiliary tank to either of said main tanks, and means to rapidly dump fuel from said interior auxiliary tank overboard.

14. Apparatus according to claim 13 and including means to isolate said in-flight pressure refueling fprobe from said fuel system when fuel is being fed to said tanks through said ground pressure refueling adapter.

15. Apparatus according to claim 14 and wherein said fuel shut-off means includes a shut-off valve and a level sensor in each of said main tanks, said shut-off valve and level sensor constructed and connected so that said shut-off valve shuts off fuel ow into said main tank when Vsaid level sensor indicates that the fuel level in said tank has reached a preselected level.

16. A helicopter having:

(a) afuselage,

(b) a sponsor projecting laterally from the port and starboard sides of said fuselage,

(c) a lift rotor supported for rotation above said fuselage,

(d) a gas turbine powerplant having an air compressor section and mounted in said fuselage to drive said lift rotor,

(e) a fuel system carried by said fuselage to provide fuel to said powerplant and including:

(l) an extendable and retractable in-ilight pressure refueling probe actuated by air pressure from said engine and dimensioned so that the tip thereof extends beyond the rotor area when the probe is extended,

(2) a forward main fuel tank and an after main fuel tank mounted in said fuselage,

(3) auxiliary fuel tanks connected to and supported `by said sponsons so as to be jettisonable therefrom,

(4) a tirst cross connector,

(5) a lirst conduit connecting said probe to said rst cross connector,

(6) a second conduit connecting said first cross connector to said forward main fuel tank,

(7) a second cross connector,

(t8) a third conduit connecting said rst cross connector to said second cross connector,

(9) a fourth conduit connecting said second cross connector to said aft main fuel tank,

(10) a fth conduit connecting said second cross connector to one of said auxiliary fuel tanks,

(11) a sixth conduit connecting said second cross connector to the other of said auxiliary `fuel tanks,

(12) means positioned in said second, third, fourth fifth and sixth conduits to select which of said tanks will receive fuel from said inflight pressure refueling probe,

(13) a ground pressure refueling adapter supported in said fuselage,

(14) a seventh conduit connecting said ground pressure refueling adapter to said first interconnector to permit selective refueling of said tanks through said second, third, fourth, fth and sixth conduits,

(15) means to isolate said pressure: refueling probe from said refueling system when fuel is being introduced to said tanks through said ground pressure refueling adapter,

(16) means to transfer fuel from either of said auxiliary tanks to either of said main tanks,

(17) means to transfer fuel between said auxiliary tanks,

(18) and means to selectively, rapidly dump over- 'board fuel from any of said tanks.

17. Apparatus according to claim 16 and wherein said power-plant is a gas turbine engine including an air compressor section, and further wherein said in-ight pressure refueling probe is actuated by pressurized air from said engine air compressor section, and further wherein said transfer means is responsive to pressurized air from said air compressor section.

18. Apparatus according to claim 17 and including means to refuel said main tanks by gravity, and means to vent all of said tanks during all types of refueling.

19. Apparatus according to claim 16 and including means located in conduits S and 6 to shut olf fuel 110W therethrough when said auxiliary tanks are jettisoned.

20. Apparatus according to claim 19 and including an interior auxiliary tank mounted in said fuselage at an elevation higher than said main tank, means to selectively transfer fuel by gravity from said interior auxiliary tank to either of said main tanks, and means to rapidly dump overboard the fuel from said interior auxiliary tank.

21. Apparatus according to claim 19 and including means to prevent the fuel from said interior auxiliary tank from entering said main tanks while dumped overboard.

22. Apparatus according to claim 18 and including an eighth conduit means with a unidirectional ow valve therein connecting said rst conduit to said tank vent means to thereby vent said first conduit means between said probe and said probe isolating means.

23. Apparatus according to claim 1 and including means to regulate the pressure of said compressor air used to transfer fuel from said auxiliary tanks to any other of said tanks.

24. Ahelicopter having:

(l) afuselage,

(2) a ylift: rotor supported for rotation from said fuselage,

(3) a gas turbine engine carried by said fuselage and connected to drive said lift rotor and having an air compressor section,

(4) a fuel receiving, storage and distribution system supported by said fuselage and including:

(a) at least two fuel tanks,

(1b) an in-flight pressure refueling probe and actuator assembly mounted to be operated between an extended and retracted probe position and dimensioned so that the probe extends beyond the lift rotor when in its extended position,

(c) and means connecting said air compressor section of said engine to Isaid probe and actuator assembly to power said probe and actuator assembly between its extended and retracted positions.

Z5. Apparatus according to claim 24 wherein said probe and actuator assembly are a composite unit and including quick attach-detach means to connect said probe and actuator assembly to and disconnect said probe and actuator assembly from the exterior of said fuselage` 26. Apparatus according to claim 24| and including means connecting said air compressor section of said engine to one of said fuel tanks to permit selective pressurization of -said tank to transfer fuel therefrom.

27. Apparatus according to claim 26 ywherein said tank which is connected to said air compressor section of said engine is mounted externally of said fuselage, and including means to jettison said tank.

being rapidly References Cited UNITED STATES PATENTS 2,516,150 7/1950 Samir-.m 244-135XR 2,580,467 1/1952 samiran. 2,771,090 11/1956 MacGregor et a1. 244-135 XR U.S. C1. X.R. 244- UNITED STATES PATENT OFFICE CERTIFICATE OF CORRECTION Patnt No. 3,432,121 March 11, 1969 Everett W. Delaney It is certified that error appears lin the above identified patent and that said Letters Patent are hereby corrected as shown below:

Column 16, line 34, after "transfer" insert fuel Signed and sealed this 31st day of March 1970.

(SEAL) Attest;

Edward M. Fletcher, J r.

Attesting Officer Commissioner of Patents WILLIAM E. SCHUYLER, JR. 

